Cooled rocket nozzle design



Oct. 20, 1964 J. c. PRossER cooLED ROCKET NozzLE DESIGN Filed NOV. 8, 1962 United States Patent() 3,153,320 COLED RGCKET NOZZLE DESIGN Joe Calvin Prosser, Laeloga, Ind., assigner to General Motors Corporation, Detroit, Mich., a corporation of Delaware Filed Nov. 8, 1962, Ser. No. 236,223

4 Claims. (Cl. 64I-35.6)

This invention relates to a rocket nozzle construction and more particularly to a nozzle constructionfacilitating transpiration cooling. v Y v Rocket nozzles frequently operate at temperatures considerably higher than the melting temperatures of the material with which the nozzle is lined.v Some form of cooling system must, therefore, be provided to prevent the liner from burning out before the termination of the operation of the nozzle. Present and past attempts to build satisfactory high perfomance cooled rocket nozzles have had limited success due to the basic problems in refractory metals, formabilityV and V4brittleness, porous material characteristics, pressure distribution in the nozzle and nozzle weights. Many designs which have evolved are quite complex or heavy and require fabrication procedures which exceed the state of the art for materials which can be used. In addition, many configurations do not have adequate strength and shock resistance to withstand the initial firing shock.

The object of this invention is to provide a relatively lightweight rocket nozzle which is inherently rugged and will be capable of operating at motive uid temperatures in excess of the melting point of its refractory metal liner by virtue of coolant provisions in the design. A further object of the invention is to construct a nozzle containing porous members contacted by a coolant for effusion of the coolant through the porous members. It is still a further object of this invention to' circumferentially surround the throat section of a porous convergentdivergent type nozzle with cooling medium manifolds to cool the higher temperature portions of the nozzle by the effusion of the coolant through the porous material.

Other objects, advantages and features of the invention will become apparent upon reference to the succeeding can be constructed by any other common 'techniques v'disc ofirefractory metal 20. It has a flange portion 22 spun or drawn on the inside margin to form a relativelyV short segment of thenzzle wall. The-thickness and choice of refractory metal can be varied to suit the operating; environmentof the nozzle.- The flange anglel 24 maybe varied With each successive ring to form the desired contour of the nozzlesurface. Y As in .this case,

- able spacing and flange length is shown on the drawingsA a convergent-divergent configuration is attained. The length of the fiange section will-depend upon the relative formability of the material and the pressure distributionV desired -in each coolant space between the Washers. Varito illustrate this feature. f

Located coaxially adjacent to and between the annular flanged discs, and extending radially inward to the flanges,

. are` porous spacer rings 26. These porous spacer rings 26 form the medium through which ows the coolant madetailed description of the preferred embodiment of the away and in section, of a nozzle embodying the invention; v

FIGURE 2 is an enlarged sectional view of the nozzle construction as shown in FIGURE 1; and

FIGURE 3 is a perspective view of a single anged y annular disc construction unit. 'n

In general, the invention relates to a convergent-divergent nozzle for high-temperature gases, cooled by transpiration-type cooling means, having a throat portion made of a number of axially ystacked thin disc-like sections. Each section includes an annulardisc having a flange portion on its inner margin, said flange forming a segment of the throat surface of the nozzle, and a porous spacer ring abutting the annular flanged disc and extending radially inward to the flanges. The porous spacer rings lie between adjacent discs and provide the medium through which flows the coolant material.

More specifically, FIGURE 1 shows a fixed nozzle 10 of the convergent-divergent type. It has a diverging conical exhaust gas exit port 12, annular throat or venturi portion 14, and a converging conical inlet portion 16 18 may constitute a portion of a rocket engine chamber or may be the aft end of any other suitable reaction motor duct. The throat and exit portions may be formed integral with the converging portions, or may be secured thereto by any suitable means.

Casing 1S is the nozzle throat supporting structure and v connected to the aft end of annular casing 18. Casing` terial. Depending upon operating conditions the porousV spacerrings may be made from a high temperature porous. ceramic, vporous graphite, or a refractory metal honeycomb.' vThe outside diameter and relative porosity of flow and distribution. In addition, a secondary, reserve type of coolant may be incorporated in the porous spacers rand be released by the heat of operation near the end of the ring time. The -primary function of the spacer 26 then jis to maintain the axial position of the refractory washers, 'support pressure loads, and aid coolant dis-V tribution. E

Since the ange portions 22 of the annular discs 20are less extensive in the direction axially of the nozzle than the width of the porous spacer rings 26, small circumferential passages 28 are formed through whichows the coolant from the porous spacer rings 26 out onto the v throat surface 14. As a result, the circumferential slots 28 are formed thereby affording the throat surface with a good distribution of coolant. The width and configuration of these slots may be varied to aid coolant flow control and pressure distribution in the coolant chamber. As is seen in FIGURE 2', the width of the porous spacer rings may'bervaried thereby permitting. a greater concentration of circumferential coolant openings where desired. f

As is shown in FIGURE 2, 'an inlet manifold 32 is provided for the coolant. tapered manifold chambers 34. YThe number of inlet manifolds 32 and chambers 34 may be'selected to suit i the amount of coolant required as determined by the operating temperature-.and time. The `manifold chambers 34 taper toward the'exit ofthe nozzleto aid controll of,l coolant flow by pressure drop. Someof theA refractory V metal vannulandiscs `20could extend out --far-ther intof, a chambers 34 to form Vvarious size orices `to further con-v rtrol coolant ow ifdesired.

In operation, the primary coolant, which wouldbe av suitable'gas or liquid, will flow through the inlet manifolds 32 into the manifold chambers 34, from whichit will effus'e through the porous spacer rings 26 towards the throat surface 14. When it'reaches the flange portion 22 it will pass through thefopenings 28.and flow out onto the nozzle throat surface14. This, then is a novel WayV to provide transpiration coolantfto a rocket nozzle throat surface.

This invention embodies many advantages over previous configurations. The relatively large percentage of light-Weight porous materialin the nozzle will-result in a very light-weight nozzle. 'TheA invention uses forms of refractory metal which are readily available.` Therela- Patented Oct. 20, 1964 This inlet manifold 32Y leads intok tively weak and thermally shock sensitive porous divider material is divided into less shock sensitive sizes and cracking of the spacer material would be of little consequence. The flexibility of the design will enable the nozzle to use coolant elfectively by adjusting the flow to suit requirements. The thin refractory metal flange will facilitate heat transfer. The division of the liner into small segments will aid thermal expansion problems. The

porosity of the spacers can be varied at will to controlV the coolant ow. Quality control and fabrication will be good because each component can be inspected individually prior to assembly. The thickness of the refractory metal can be Varied from disc to disc to compensate for erosion and thereby produce a minimum weight design. And finally, the nozzle is inherently rugged and will function as a heat sink after exhaustion of coolant for added ring time.

While the invention has been illustrated for use in connection with a rocket motor casing, it will be clear that it would have use in any installation other than those illustrated, and that many modifications and changes may be made thereto without departing from the scope of the invention.

I claim:

1. A nozzle, cooled by a transpiration-type cooling means, for high temperature gases comprising, in combination, a coaxial stack of a plurality of annular discs each having a flange portion on its inside margin, said ange forming a segment of the throat surface of the nozzle, and porous spacer rings located coaxial with, ad-

jacent to, and between said annular flanged discs and extending radially inward to said flanges, -said porous spacer rings forming the medium through which the coolant material ilows. Y

2. A nozzle as described in claim l with said flange portion slightly less extensive in the direction axially of the nozzle than the width of the porous spacer rings thereby forming a narrow coolant passage around the circumference of the throat through which the coolant flows from said porous spacer rings out onto the throat surface.

3. A nozzle as described in claim l with said ange portions of the annular rings having a Varying ange angle Reerences Cited in the le of this patent UNITED STATES PATENTS 3,026,806 lRunton et al. --.1 Mar. 27, 1962 3,048,972 Barlow Aug. 14, 1962 3,069,847 Vest Dec. 25, 1962 3,103,885 McLauchlan ..-a Sept. 17,' 1963 

1. A NOZZLE, COOLED BY A TRANSPIRATION-TYPE COOLING MEANS, FOR HIGH TEMPERATURE GASES COMPRISING, IN COMBINATION, A COAXIAL STACK OF A PLURALITY OF ANNULAR DISCS EACH HAVING A FLANGE PORTION ON ITS INSIDE MARGIN, SAID FLANGE FORMING A SEGMENT OF THE THROAT SURFACE OF THE NOZZLE, AND POROUS SPACER RINGS LOCATED COAXIAL WITH, ADJACENT TO, AND BETWEEN SAID ANNULAR FLANGED DISCS AND EXTENDING RADIALLY INWARD TO SAID FLANGES, SAID POROUS SPACER RINGS FORMING THE MEDIUM THROUGH WHICH THE COOLANT MATERIAL FLOWS. 